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NASA NTRS1985

Advanced thermal barrier coating systems

Current state-of-the-art thermal barrier coating (TBC) systems consist of partially stabilized zirconia coatings plasma sprayed over a MCrAlY bond coat. Although these systems have excellent thermal shock properties, they have shown themselves to be deficient for a number of diesel and aircraft applications. Two ternary ceramic plasma coatings are discussed with respect to their possible use in TBC systems. Zirconia-ceria-yttria (ZCY) coatings were developed with low thermal conductivities, good thermal shock resistance and improved resistance to vanadium containing environments, when compared to the baseline yttria stabilized zirconia (YSZ) coatings. In addition, dense zirconia-titania-yttria (ZTY) coatings were developed with particle erosion resistance exceeding conventional stabilized zirconia coatings. Both coatings were evaluated in conjunction with a NiCr-Al-Co-Y2O3 bond coat. Also, multilayer or hybrid coatings consisting of the bond coat with subsequent coatings of zirconia-ceria-yttria and zirconia-titania-yttria were evaluated. These coatings combine the enhanced performance characteristics of ZCY with the improved erosion resistance of ZTY coatings. Improvement in the erosion resistance of the TBC system should result in a more consistent delta T gradient during service. Economically, this may also translate into increased component life simply because the coating lasts longer.

Dorfman, M. R.↗

NASA NTRS2001

Thermal Processing Effects on the Adhesive Strength of PS304 High Temperature Solid Lubricant Coatings

In this paper the effects of post deposition heat treatments on the cohesive and adhesive strength properties of PS304, a plasma sprayed nickel-chrome based, high temperature solid lubricant coating deposited on stainless steel, are studied. Plasma spray deposited coating samples were exposed in air at temperatures from 432 to 650 C for up to 500 hr to promote residual stress relief, enhance particle to particle bonding and increase coating to substrate bond strength. Coating pull-off strength was measured using a commercial adhesion tester that utilizes 13 mm diameter aluminum pull studs attached to the coating surface with epoxy. Pull off force was automatically recorded and converted to coating pull off strength. As deposited coating samples were also tested as a baseline. The as-deposited (untreated) samples either delaminated at the coating-substrate interface or failed internally (cohesive failure) at about 17 MPa. Samples heat treated at temperatures above 540 C for 100 hr or at 600 C or above for more than 24 hr exhibited strengths above 31 MPa, nearly a two fold increase. Coating failure occurred inside the body of the coating (cohesive failure) for nearly all of the heat-treated samples and only occasionally at the coating substrate interface (adhesive failure). Metallographic analyses of heat-treated coatings indicate that the Nickel-Chromium binder in the PS304 appears to have segregated into two phases, a high nickel matrix phase and a high chromium precipitated phase. Analysis of the precipitates indicates the presence of silicon, a constituent of a flow enhancing additive in the commercial NiCr powder. The exact nature and structure of the precipitate phase is not known. This microstructural change is believed to be partially responsible for the coating strength increase. Diffusion bonding between particles may also be playing a role. Increasing the heat treatment temperature, exposure time or both accelerate the heat treatment process. Preliminary measurements indicate that the heat treatment also results in a one time, permanent coating thickness increase of about 3%. Based upon these results, the incorporation of a heat treatment prior to final finishing has been incorporated in the application process of this coating technology.

DellaCorte, Christopher↗

NASA NTRSDate not supplied

Conductive Tether Coating for Electrodynamic Tethers

The Propulsive Small Expendable Deployer System (ProSEDS), which is an on-orbit demonstration of the propulsion capabilities of electrodynamic tethers in space, is a secondary payload on a Delta 11 unmanned expendable booster. The ProSEDS tether consists of a 5 km bare electrodynamic tether and a 1 0-km non-conductive leader tether. Near the Delta 11, 160 m of the conductive tether is insulated to prevent plasma electron collection from the plasma contactor and for other science requirements. The remainder of the 5-km conductive tether is coated with a new conductive coating to collect plasma electrons. A bare metal tether easily collects electrons from the plasma, but thermal concerns preclude this design. A highly emissive conductive polymer developed by Triton Systems, Inc. has been optimized for both conductivity and thermo-optical properties. The current design for the ProSEDS conductive tether is seven individually coated strands of 28 AWG aluminum wire, coated with an atomic oxygen-resistant conductive polymer composed of a mixture of COR (Colorless Oxygen Resistant) and polyanaline (PANI) known as C-COR (Conductive-Colorless Oxygen Resistant). The conductive-coated wire strands are cold-welded to individually coated strands of the insulated tether. The insulated tether is coated with 1 mil of polyimide and an atomic oxygen resistant polymer TOR-BP. The insulated tether must stand off the entire voltage of the tether (1 200 V) at various times during the mission. All seven wires are twisted around a Kevlar-29 core using the Hi-wire design. Extensive testing has been performed at the Marshall Space Flight Center to qualify both the conductive coating and insulating coating for use on the ProSEDS tether. The conductive coating has been exposed to a plasma to verify the coatings ability to collect electrons from the space plasma from 0 to 1500 V, and to verify the coatings ability to collect electrons after atomic oxygen exposure. The insulated coating has been tested to verify its dielectric strength both in a vacuum and air. Thermal optical properties on both coatings have been measured before and after atomic oxygen testing. The results of these tests will be presented.

Jason A Vaughn↗

NASA NTRS2000

Erosion Coatings for High-Temperature Polymer Composites: A Collaborative Project With Allison Advanced Development Company

The advantages of replacing metals in aircraft turbine engines with high-temperature polymer matrix composites (PMC's) include weight savings accompanied by strength improvements, reduced part count, and lower manufacturing costs. Successfully integrating high-temperature PMC's into turbine engines requires several long-term characteristics. Resistance to surface erosion is one rarely reported property of PMC's in engine applications because PMC's are generally softer than metals and their erosion resistance suffers. Airflow rates in stationary turbine engine components typically exceed 2.3 kg/sec at elevated temperatures and pressures. In engine applications, as shown in the following photos, the survivability of PMC components is clearly a concern, especially when engine and component life-cycle requirements become longer. Although very few publications regarding the performance of erosion coatings on PMC's are available particularly in high-temperature applications the use of erosion-resistant coatings to significantly reduce wear on metallic substrates is well documented. In this study initiated by the NASA Glenn Research Center at Lewis Field, a low-cost (less than $140/kg) graphite-fiber-reinforced T650 35/PMR 15 sheet-molding compound was investigated with various coatings. This sheet-molding compound has been compression molded into many structurally complicated components, such as shrouds for gas turbine inlet housings and gearboxes. Erosion coatings developed for PMC s in this study consisted of a two-layered system: a bondcoat sprayed onto a cleaned PMC surface, followed by an erosion-resistant, hard topcoat sprayed onto the bondcoat as shown in following photomicrograph. Six erosion coating systems were evaluated for their ability to withstand harsh thermal cycles, erosion resistance (ASTM G76 83 "Standard Practice for Conducting Erosion Tests by Solid Particle Impingement Using Gas Jets") using Al2O3, and adhesion to the graphite fiber polyimide composite (ASTM D 4541 95 "Pull Off Strength of Coatings"). Glenn and Allison Advanced Development Company collaborated to optimize erosion coatings for gas turbine fan and compressor applications. All the coating systems survived aggressive thermal cycling without spalling. During erosion tests (see the final photo), the most promising coating systems tested had Cr3C2-NiCr and WC-Co as the hard topcoats. In all cases, these coating systems performed significantly better than that with a TiN hard topcoat. When material depth (thickness) loss is considered, the Cr3C2-NiCr and WC-Co coating systems provided, on average, an erosion resistance 8.5 times greater than that for the uncoated PMR 15/T650 35 composite. Similarly, Cr3C2-NiCr and WC-Co coating systems adhered to the PMC substrate during tensile tests significantly better than systems containing a TiN topcoat. Differences in topcoats of Cr3C2-NiCr and WC-Co were determined by considering issues such as cost and environmental impact. The preferred erosion-resistant coating system for PMR 15/T650 35 has WC-Co as the hard topcoat. This system provides the following benefits in comparison to the coating system with Cr3C2-NiCr topcoat: lower powder material cost (15 to 20 percent), environmentally friendly materials (Cr3C2-NiCr is hazardous), and higher deposition yield (10 to 15 percent), which results in less waste.

Sutter, James K.↗

NASA NTRS1990

Development of Low Conductivity and Ultra High Temperature Ceramic Coatings Using A High-Heat-Flux Testing Approach

The development of low conductivity, robust thermal and environmental barrier coatings requires advanced testing techniques that can accurately and effectively evaluate coating thermal conductivity and cyclic resistance at very high surface temperatures (up to 17OOOC) under large thermal gradients. In this study, a laser high-heat-flux test approach is established for evaluating advanced low conductivity, ultra-high temperature ceramic thermal and environmental barrier coatings under the NASA Ultra Efficient Engine Technology (UEET) program. The test approach emphasizes the real-time monitoring and assessment of the coating thermal conductivity: the initial conductivity rise under a steady-state high temperature thermal gradient test due to coating sintering, and the later coating conductivity reduction under a subsequent cyclic thermal gradient test due to coating cracking/delamination. The coating system is then evaluated based on the damage accumulations and failure after the combined steady-state and cyclic thermal gradient tests. The lattice and radiation thermal conductivity of advanced ceramic coatings can also be evaluated using laser heat-flux techniques. The coating external radiation resistance is assessed based on the measured specimen temperature response under a laser heated intense radiation flux source. The coating internal radiation contribution is investigated based on the measured apparent coating conductivity increases with the coating surface test temperature under large thermal gradient test conditions. Since an increased radiation contribution is observed at these very high surface test temperatures, by varying the laser heat-flux and coating average test temperature, the complex relation between the lattice and radiation conductivity as a function of surface and interface test temperature is derived.

Zhu, Dongming↗

NASA NTRS2008

Final Report on Portable Laser Coating Removal Systems Field Demonstrations and Testing

Processes currently used throughout the National Aeronautics and Space Administration (NASA) to remove corrosion and coatings from structures, ground service equipment and small components results in waste streams consisting of toxic chemicals, spent media blast materials, and waste water. When chemicals are used in these processes they are typically high in volatile organic compounds (VOC) and are considered hazardous air pollutants (HAP). When blast media is used, the volume of hazardous waste generated is increased significantly. Many of the coatings historically used within NASA contain toxic metals such as hexavalent chromium, and lead. These materials are highly regulated and restrictions on worker exposure continue to increase. Most recently the EPA reduced the permissible exposure limit (PEL) for hexavalent chromium. The new standard lowers OSHA's PEL for hexavalent chromium from 52 to 5 micrograms of Cr(V1) per cubic meter of air as an 8-hour time-weighted average. Hexavalent chromium is found in the pretreatment and primer coatings used within the Shuttle Program. In response to the need to continue to protect assets within the agency and the growing concern over these new regulations, NASA is researching different ways to continue the required maintenance of both facility and flight equipment in a safe, efficient and environmentally preferable manner. The use of laser energy to remove prepare surfaces for a variety of processes, such as corrosion and coating removal, weld preparation and non destructive evaluation is a relatively new technology that has shown itself to be environmentally preferable and in many cases less labor intensive than currently used removal methods. The development of a Portable Laser Coating Removal System (PLCRS) started as the goal of a Joint Group on Pollution Prevention (JG-PP) project, led by the Air Force, where several types of lasers in several configurations were thoroughly evaluated. Following this project, NASA decided to evaluate the best performers on processes and coatings specific to the agency. Laser systems used during this project were all of a similar design, most of which had integrated vacuum systems in order to collect materials removed from substrate surfaces during operation. Due to the fact that the technology lends itself to a bide variety of processes, several site demonstrations were organized in order to allow for greater evaluation of the laser systems across NASA. The project consisted of an introductory demonstration and a more in-depth evaluation at Wright-Patterson Air Force Base. Additionally, field demonstrations occurred at Glenn Research Center and Kennedy Space Center. During these demonstrations several NASA specific applications were evaluated, including the removal of coatings within Orbiter tile cavities and Teflon from Space Shuttle Main Engine gaskets, removal of heavy grease from Solid Rocket Booster components and the removal of coatings on weld lines for Shuttle and general ground service equipment for non destructive evaluation (NDE). In addition, several general industry applications such as corrosion removal, structural coating removal, weld-line preparation and surface cleaning were evaluated. This included removal of coatings and corrosion from surfaces containing lead-based coatings and applications similar to launch-structure maintenance and Crawler maintenance. During the project lifecycle, an attempt was made to answer process specific concerns and questions as they arose. Some of these initially unexpected questions concerned the effects lasers might have on substrates used on flight equipment including strength, surface re-melting, substrate temperature and corrosion resistance effects. Additionally a concern was PPE required for operating such a system including eye, breathing and hearing protection. Most of these questions although not initially planned, were fully explored as a part of this project. Generally the results from tesng were very positive. Corrosion was effectively removed from steel, but less successfully from aluminum alloys. Coatings were able to be removed, with varying results, generally dark, matte and thin coatings were easier to remove. Steel and aluminum panels were able to be cleaned for welding, with no known deleterious effects and weld-lines were able to have coatings removed in critical areas for NDE while saving time as compared to other methods.

Rothgeb, Matthew J.↗

NASA NTRS2008

In-Service Evaluation of HVOF Coated Main Landing Gear on Navy P-3 Aircraft

Due to the environmental and health concerns with Electroplated Hard Chrome (EHC), the Hard Chrome Alternatives Team (HCAT) has been working to provide an alternative wear coating for EHC. The US Navy selected Tungsten-Carbide Cobalt (WC- 17Co) High Velocity Oxy-Fuel (HVOF) thermal spray coating for this purpose and completed service evaluations on select aircraft components to support the HCAT charter in identifying an alternative wear coating for chrome plating. Other benefits of WC-Co thermal spray coatings over EHC are enhanced corrosion resistance, improved durability, and exceptional wear properties. As part of the HCAT charter and to evaluate HVOF coatings on operational Navy components, the P-3 aircraft was selected for a service evaluation to determine the coating durability as compared to chrome plating. In April 1999, a VP-30 P-3 aircraft was outfitted with a right-hand Main Landing Gear (MLG) shock strut coated with WCCo HYOF thermal spray applied to the piston barrel and four axle journals. The HVOF coating on the piston barrel and axle journals was applied by Southwest United Industries, Inc. This HVOF coated strut assembly has since completed 6,378 landings. Teardown analysis .for this WC-Co HVOF coated MLG asset is significant in assessing the durability of this wear coating in service relative to EHC and to substantiate Life Cycle Cost (LCC) data to support a retrograde transition from EHC to HVOF thermal spray coatings. Findings from this teardown analysis may also benefit future transitions to HVOF thermal spray coatings by identifying enhancements to finishing techniques, mating bearing and liner material improvements, improved seal materials, and improvements in HVOF coating selection.

Devereaux, jon L.↗

NASA NTRS2016

Environmental Barrier Coatings for Ceramic Matrix Composites - An Overview

Ceramic Matrix Composites (CMCs) are increasingly being considered as structural materials for advanced power generation equipment. Broadly speaking the two classes of materials are oxide-based CMCs and non-oxide based CMCs. The non-oxide CMCs are primarily silicon-based. Under conditions prevalent in the gas turbine hot section the water vapor formed in the combustion of gaseous or liquid hydrocarbons reacts with the surface-SiO2 to form volatile products. Progressive surface recession of the SiC-SiC CMC component, strength loss as a result of wall thinning and chemical changes in the component occur, which leads to the loss of structural integrity and mechanical strength and becomes life limiting to the equipment in service. The solutions pursued to improve the life of SiC-SiC CMCs include the incorporation of an external barrier coating to provide surface protection to the CMC substrate. The coating system has become known as an Environmental Barrier Coating (EBC). The relevant early coatings work was focused on coatings for corrosion protection of silicon-based monolithic ceramics operating under severely corrosive conditions. The development of EBCs for gas turbine hot section components was built on the early work for silicon-based monolithics. The first generation EBC is a three-layer coating, which in its simplest configuration consists of a silicon (Si) base coat applied on top of the CMC, a barium-strontium-aluminosilicate (BSAS) surface coat resistant to water vapor attack, and a mullite-based intermediate coating layer between the Si base coat and BSAS top coat. This system can be represented as Si-Mullite-BSAS. While this baseline EBC presented a significant improvement over the uncoated SiC-SiC CMC, for the very long durations of 3-4 years or more expected for industrial operation further improvements in coating durability are desirable. Also, for very demanding applications with higher component temperatures but shorter service lives more rugged EBCs will be necessary. A second generation of EBCs incorporates rare earth silicates which have extremely favorable resistance against environmental attack and a higher temperature capability. Performance data for this class of EBCs is more limited and especially field data are not as extensive as for the first generation EBCs. Extensive laboratory, rig and engine testing, including testing of EBC coated SiC-SiC CMCs in actual field applications is in progress. The development of next generation EBCs with even higher temperature capability than the second generation EBC is also underway. This paper will discuss the current status of EBC technology and future direction based on literature survey.

Laboratory↗

NASA NTRS2017

Coatings for Oxidation and Hot Corrosion Protection of Disk Alloys

Increasing temperatures in aero gas turbines is resulting in oxidation and hot corrosion attack of turbine disks. Since disks are sensitive to low cycle fatigue (LCF), any environmental attack, and especially hot corrosion pitting, can potentially seriously degrade the life of the disk. Application of metallic coatings are one means of protecting disk alloys from this environmental attack. However, simply the presence of a metallic coating, even without environmental exposure, can degrade the LCF life of a disk alloy. Therefore, coatings must be designed which are not only resistant to oxidation and corrosion attack, but must not significantly degrade the LCF life of the alloy. Three different Ni-Cr coating compositions (29, 35.5, 45wt. Cr) were applied at two thicknesses by Plasma Enhanced Magnetron Sputtering (PEMS) to two similar Ni-based disk alloys. One coating also received a thin ZrO2 overcoat. The coated samples were also given a short oxidation exposure in a low PO2 environment to encourage chromia scale formation. Without further environmental exposure, the LCF life of the coated samples, evaluated at 760C, was less than that of uncoated samples. Hence, application of the coating alone degraded the LCF life of the disk alloy. Since shot peening is commonly employed to improve LCF life, the effect of shot peening the coated and uncoated surface was also evaluated. For all cases, shot peening improved the LCF life of the coated samples. Coated and uncoated samples were shot peened and given environmental exposures consisting of 500 hrs of oxidation followed by 50 hrs of hot corrosion, both at 760C). The high-Cr coating showed the best LCF life after the environmental exposures. Results of the LCF testing and post-test characterization of the various coatings will be presented and future research directions discussed.

LCF life↗

NASA NTRS1982

Advanced ceramic coating development for industrial/utility gas turbines

A program was conducted with the objective of developing advanced thermal barrier coating (TBC) systems. Coating application was by plasma spray. Duplex, triplex and graded coatings were tested. Coating systems incorporated both NiCrAly and CoCrAly bond coats. Four ceramic overlays were tested: ZrO2.82O3; CaO.TiO2; 2CaO.SiO2; and MgO.Al2O3. The best overall results were obtained with a CaO.TiO2 coating applied to a NiCrAly bond coat. This coating was less sensitive than the ZrO2.8Y2O3 coating to process variables and part geometry. Testing with fuels contaminated with compounds containing sulfur, phosphorus and alkali metals showed the zirconia coatings were destabilized. The calcium titanate coatings were not affected by these contaminants. However, when fuels were used containing 50 ppm of vanadium and 150 ppm of magnesium, heavy deposits were formed on the test specimens and combustor components that required frequent cleaning of the test rig. During the program Mars engine first-stage turbine blades were coated and installed for an engine cyclic endurance run with the zirconia, calcium titanate, and calcium silicate coatings. Heavy spalling developed with the calcium silicate system. The zirconia and calcium titanate systems survived the full test duration. It was concluded that these two TBC's showed potential for application in gas turbines.

Vogan, J. W.↗

NASA NTRS1985

Thermal barrier coating life prediction model

This is the first report of the first phase of a 3-year program. Its objectives are to determine the predominant modes of degradation of a plasma sprayed thermal barrier coating system, then to develop and verify life prediction models accounting for these degradation modes. The first task (Task I) is to determine the major failure mechanisms. Presently, bond coat oxidation and bond coat creep are being evaluated as potential TBC failure mechanisms. The baseline TBC system consists of an air plasma sprayed ZrO2-Y2O3 top coat, a low pressure plasma sprayed NiCrAlY bond coat, and a Rene'80 substrate. Pre-exposures in air and argon combined with thermal cycle tests in air and argon are being utilized to evaluate bond coat oxidation as a failure mechanism. Unexpectedly, the specimens pre-exposed in argon failed before the specimens pre-exposed in air in subsequent thermal cycles testing in air. Four bond coats with different creep strengths are being utilized to evaluate the effect of bond coat creep on TBC degradation. These bond coats received an aluminide overcoat prior to application of the top coat to reduce the differences in bond coat oxidation behavior. Thermal cycle testing has been initiated. Methods have been selected for measuring tensile strength, Poisson's ratio, dynamic modulus and coefficient of thermal expansion both of the bond coat and top coat layers.

Hillery, R. V.↗

NASA NTRS1988

Thermal barrier coating life prediction model development

This report describes work performed to determine the predominat modes of degradation of a plasma sprayed thermal barrier coating system and to develop and verify life prediction models accounting for these degradation modes. The primary TBC system consisted of a low pressure plasma sprayed NiCrAlY bond coat, an air plasma sprayed ZrO2-Y2O3 top coat, and a Rene' 80 substrate. The work was divided into 3 technical tasks. The primary failure mode to be addressed was loss of the zirconia layer through spalling. Experiments showed that oxidation of the bond coat is a significant contributor to coating failure. It was evident from the test results that the species of oxide scale initially formed on the bond coat plays a role in coating degradation and failure. It was also shown that elevated temperature creep of the bond coat plays a role in coating failure. An empirical model was developed for predicting the test life of specimens with selected coating, specimen, and test condition variations. In the second task, a coating life prediction model was developed based on the data from Task 1 experiments, results from thermomechanical experiments performed as part of Task 2, and finite element analyses of the TBC system during thermal cycles. The third and final task attempted to verify the validity of the model developed in Task 2. This was done by using the model to predict the test lives of several coating variations and specimen geometries, then comparing these predicted lives to experimentally determined test lives. It was found that the model correctly predicts trends, but that additional refinement is needed to accurately predict coating life.

Hillery, R. V.↗

NASA NTRS1996

Static measurement of the thickness of the ablative coating of the solid rocket boosters

The Solid Rocket Boosters (SRB's) used to launch the Space Shuttle are coated with a layer of ablative material to prevent thermal damage when they reenter the earth's atmosphere. The coating consists of a mixture of cork, glass, and resin. A new coating (Marshall Convergent Coating, MCC-2) was recently developed that is environmentally complaint. The coating must meet certain minimum thickness standards in order to protect the SRB. The coating is applied by a robot controlled nozzle that moves from the bottom to top, as the rocket part rotates on a table. Several coats are applied, building up to the desired thickness. Inspectors do a limited amount of destructive 'wet' testing. This involves an inspector inserting a rod in the wet coating and removing the rod. This results in a hole that, of course, must be patched later. The material is cured and the thickness is measured. There is no real-time feedback as the coating is being applied. Although this might seem like the best way to control thickness, the problems with 'blowback' (reflected material covering the sensor) are formidable, and have not been solved. After the thermal coating is applied, a protective top coat is applied. The SRB part is then placed in a oven and baked to harden the surface. The operations personnel then measure the thickness of the layer using the Kaman 7200 Displacement Measuring System. The probe is placed on the surface. One person (the inspector) reads the instrument, while another(the technician) records the thickness. Measurements are taken at one foot intervals. After the measurements are taken, the number of low readings is tabulated. If more than 10 percent of the points fall below the minimum value, there is a design review, and the part may be stripped of coating, and a new coating is applied. There is no other analysis.

Harrison, Harry C.↗

NASA NTRS2003

Development of Advanced Thermal and Environmental Barrier Coatings Using a High-Heat-Flux Testing Approach

The development of low conductivity, robust thermal and environmental barrier coatings requires advanced testing techniques that can accurately and effectively evaluate coating thermal conductivity and cyclic resistance at very high surface temperatures (up to 1700 C) under large thermal gradients. In this study, a laser high-heat-flux test approach is established for evaluating advanced low conductivity, high temperature capability thermal and environmental barrier coatings under the NASA Ultra Efficient Engine Technology (UEET) program. The test approach emphasizes the real-time monitoring and assessment of the coating thermal conductivity, which initially rises under the steady-state high temperature thermal gradient test due to coating sintering, and later drops under the cyclic thermal gradient test due to coating cracking/delamination. The coating system is then evaluated based on damage accumulation and failure after the combined steady-state and cyclic thermal gradient tests. The lattice and radiation thermal conductivity of advanced ceramic coatings can also be evaluated using laser heat-flux techniques. The external radiation resistance of the coating is assessed based on the measured specimen temperature response under a laser- heated intense radiation-flux source. The coating internal radiation contribution is investigated based on the measured apparent coating conductivity increases with the coating surface test temperature under large thermal gradient test conditions. Since an increased radiation contribution is observed at these very high surface test temperatures, by varying the laser heat-flux and coating average test temperature, the complex relation between the lattice and radiation conductivity as a function of surface and interface test temperature may be derived.

Zhu, Dongming↗

NASA NTRS2015

Anti-Adhesion Elastomer Seal Coatings for Ultraviolet and Atomic Oxygen Protection

Radiation blocking sunscreen coatings have been developed for the protection of elastomer seals used in low-Earth-orbit (LEO). The coatings protect the seals from ultraviolet (UV) radiation and atomic oxygen (AO) damage. The coatings were developed for use on NASA docking seals. Docking seal damage from the UV and AO present in LEO can constrain mission time-line, flight mode options, and increases risk. A low level of adhesion is also required for docking seals so undocking push-off forces can be low. The coatings presented also mitigate this unwanted adhesion. Greases with low collected volatile condensable materials (CVCM) and low total mass loss (TML) were mixed with slippery and/or UV blocking powders to create the protective coatings. Coatings were applied at rates up to 2 milligrams per square centimeter. Coated seals were exposed to AO and UV in the NUV (near-UV) and UV-C wavelength ranges (300 to 400 nanometers and 254 nanometers, respectively). Ground based ashers were used to simulate the AO of space. The Sun's UV energy was mimicked assuming a nose forward flight mode, resulting in an exposure rate of 2.5 megajoules per square meter per day. Exposures between 0 and 147 megajoules per square meter (UV-C) and 245 megajoules per square meter (NUV) were accomplished. The protective coatings were durable, providing protection from UV after a simulated docking and undocking cycle. The level of protection begins to decline at coverage rates less than 0.9 milligrams per square centimeter. The leakage of seals coated with Braycote plus 20 percent Z-cote ZnO sunscreen increased by a factor of 40 after moderate AO exposure; indicating that this coating might not be suitable due to AO intolerance. Seals coated with DC-7-16.4 percent Z-cote ZnO sunscreen were not significantly affected by combined doses of 2 x 10 (sup 21) atoms per square AO with 73 megajoules per square meter UV-C. Unprotected seals were significantly damaged at UV-C exposures of 0.3 megajoules per square meter and DC-7-16.4 percent Z-cote coated seals were undamaged at all exposures up to the limits tested thus far which were 147 megajoules per square meter UV-C and 245 megajoules per square meter NUV. The coatings decreased adhesion sufficiently for docking seals at temperatures equal to or greater than -8 degrees Centigrade thus offer a simple and inexpensive way to mitigate adhesion.

adhesion↗

NASA NTRS2018

Reflective Coatings for the Future X-Ray Mirror Substrates

We present the development of the reflective coating by magnetron sputtering deposition onto precisely-fabricated thin X-ray mirrors. Our goal is to remove distortion induced by the coating and then keep their surface pro les. We first addressed the uniform coating to minimize the distortion by introducing a mask to control the spatial distribution of the coating thickness. The uniformity was finally achieved within 1%. We next tried a platinum single-layer coating on a glass substrate with a dimension of 200 mm 125 mm. The distortion caused by the frontside coating with a thickness of 320 A was found to be at most 1 m, smaller than the previous results obtained from the non-uniform coating. We then carried out the platinum coating with the same amount of the thickness on the backside surface of the glass substrate. The surface pro le of the glass substrate was fully recovered, indicating that the residual stress was successfully balanced by the backside coating. Furthermore, we tried to an iridium single-layer coating with a thickness of 150 Aon the silicon mirrors. The frontside coating caused the degradation of the imaging quality by 7:5 arcsec in half-power width. However, the backside coating with the same amount of the thickness reduced this degradation to be 3:4 arcsec. Finally, an additional backside coating with a thickness of 100 A and the annealing to relax the residual stress were found to eliminate the distortion completely; the final degradation of the imaging quality was only 0:4 arcsec.

Mori, Hideyuki↗

NASA NTRS1985

Rocket thrust chamber thermal barrier coatings

A research program was conducted to generate data and develop analytical techniques to predict the performance and reliability of ceramic thermal barrier coatings in high heat flux environments. A finite element model was used to analyze the thermomechanical behavior of coating systems in rocket thrust chambers. Candidate coating systems (using a copper substrate, NiCrAlY bond coat and ZrO2.8Y2O3 ceramic overcoat) were selected for detailed study based on photomicrographic evaluations of experimental test specimens. The effects of plasma spray application parameters on the material properties of these coatings were measured and the effects on coating performance evaluated using the finite element model. Coating design curves which define acceptable operating envelopes for seleted coating systems were constructed based on temperature and strain limitations. Spray gun power levels was found to have the most significant effect on coating structure. Three coating systems were selected for study using different power levels. Thermal conductivity, strain tolerance, density, and residual stress were measured for these coatings. Analyses indicated that extremely thin coatings ( 0.02 mm) are required to accommodate the high heat flux of a rocket thrust chamber and ensure structural integrity.

Batakis, A. P.↗

NASA NTRS1995

Thermal conductivity of zirconia thermal barrier coatings

Thermal barrier coatings (TBC's) applied to the hot gas components of turbine engines lead to enhanced fuel efficiency and component reliability. Understanding the mechanisms which control the thermal transport behavior of the TBC's is of primary importance. Physical vapor description (PVD) and plasma spraying (PS) are the two most commonly used coating techniques. These techniques produce coatings with unique microstructures which control their performance and stability. The PS coatings were applied with either standard power or hollow sphere particles. The hollow sphere particles yielded a lower density and lower thermal conductivity coating. The thermal conductivity of both fully and partially stabilized zirconia, before and after thermal aging, will be compared. The thermal conductivity of the coatings permanently increase upon being exposed to high temperatures. These increases are attributed to microstructural changes within the coatings. Sintering of the as fabricated plasma sprayed lamellar structure is observed by scanning electron microscopy of coatings isothermally heat treated at temperatures greater than 1100 C. During this sintering process the planar porosity between lamella is converted to a series of small spherical pores. The change in pore morphology is the primary reason for the observed increase in thermal conductivity. This increase in thermal conductivity can be modeled using a relationship which depends on both the temperature and time of exposure. Although the PVD coatings are less susceptible to thermal aging effects, preliminary results suggest that they have a higher thermal conductivity than PS coatings, both before and after thermal aging. The increases in thermal conductivity due to thermal aging for partially stabilized plasma sprayed zirconia have been found to be less than for fully stabilized plasma sprayed zirconia coatings. The high temperature thermal diffusivity data indicates that if these coatings reach a temperature above 1100 C during operation, they will begin to lose their effectiveness as a thermal barrier.

Dinwiddie, R. B.↗